1. Field of the Invention
The present invention relates to an air fuel mixer for the combustor of a gas turbine engine, and, more particularly, to an air fuel mixer for the combustor of a high pressure ratio gas turbine engine which uniformly mixes fuel and air so as to reduce NOx formed by the ignition of the fuel/air mixture.
2. Description of Related Art
Air pollution concerns worldwide have led to stricter emissions standards requiring significant reductions in gas turbine pollutant emissions, especially for industrial and power generation applications. Nitrous Oxide (NOx), which is a precursor to atmospheric pollution, is generally formed in the high temperature regions of the gas turbine combustor by direct oxidation of atmospheric nitrogen with oxygen. Reductions in gas turbine emissions of NOx have been obtained by the reduction of flame temperatures in the combustor, such as through the injection of high purity water or steam in the combustor. Additionally, exhaust gas emissions have been reduced through measures such as selective catalytic reduction. While both the wet techniques (water/steam injection) and selective catalytic reduction have proven themselves in the field, both of these techniques require extensive use of ancillary equipment. Obviously, this drives the cost of energy production higher. Other techniques for the reduction of gas turbine emissions include "rich burn, quick quench, lean burn" and "lean premix" combustion, where the fuel is burned at a lower temperature.
In a typical aero-derivative industrial gas turbine engine, fuel is burned in an annular combustor. The fuel is metered and injected into the combustor by means of multiple nozzles into a venturi along with combustion air having a designated amount of swirl. No particular care has been exercised in the prior art, however, in the design of the nozzle, the venturi or the dome end of the combustor to mix the fuel and air uniformly to reduce the flame temperatures. Accordingly, non-uniformity of the air/fuel mixture causes the flame to be locally hotter, leading to significantly enhanced production of NOx.
In the typical aircraft gas turbine engine, flame stability and variable cycle operation of the engine dominate combustor design requirements. This has in general resulted in combustor designs with the combustion at the dome end of the combustor proceeding at the highest possible temperatures at stoichiometric conditions. This, in turn, leads to large quantities of NOx to be formed in such gas turbine combustors since it has been of secondary importance.
While premixing ducts in the prior art have been utilized in lean burning designs, they have been found to be unsatisfactory due to flashback and auto-ignition considerations for modern gas turbine applications. Flashback involves the flame of the combustor being drawn back into the mixing section, which is most often caused by a backflow from the combustor due to compressor instability and transient flows. Auto-ignition of the fuel/air mixture can occur within the premixing duct if the velocity of the air flow is not fast enough, i.e., where there is a local region of high residence time. Flashback and auto-ignition have become serious considerations in the design of mixers for aero-derivative engines due to increased pressure ratios and operating temperatures. Since one desired application of the present invention is for the LM6000 gas turbine engine, which is the aero-derivative of General Electric's CF6-80C2 engine, these considerations are of primary significance.
While the effects of counter-rotating swirl have been studied (.e.g., "Effectiveness of Mixing Coaxial Flows Swirled in Opposite Directions," by A. Sviridenkov, V. Tret'yakov, and V. Yagodkin; "Distribution of Velocity Pulsations in a Channel with Mixing of Oppositely Swirled Streams," by A. Sviridenkov and V. Tret'yakov; and "Reactive Mixing in Swirling Flows," by W. Cheng), they have not been utilized with fuel injection techniques that uniformly premix the fuel and air prior to combustion. Likewise, fuel nozzles and injectors which inject fuel into an air flow for premixing, such as the radial fuel spokes in "Experimental Evaluation of a Low Emissions, Variable Geometry, Small Gas Turbine Combustor," by K. O. Smith, M. H. Smaii, and H. K. Mak and the fuel injector having a conical tip in U.S. Pat. No. 4,653,278 to Vinson et al, neither combine with the intense shear region provided by counter-rotating swirlers.
An air fuel mixer is disclosed in U.S. Pat. No. 5,165,241, also owned by the assignee of the present invention, which includes a mixing duct, a set of inner and outer counter-rotating swirlers at the upstream end of the mixing duct, and a fuel nozzle located axially along and forming a centerbody of the mixing duct. While high pressure air from a compressor is injected into the mixing duct through the swirlers to form an intense shear region, the fuel is injected into the mixing duct from holes formed in the fuel nozzle. This configuration for fuel injection is adequate for low pressure ratio engines, but it has been found lacking for high pressure ratio engines.
Accordingly, a primary objective of the present invention is to provide an air fuel mixer for a high pressure ratio aero-derivative gas turbine engine which avoids the problems of auto-ignition and flashback.
Yet another objective of the present invention is to provide an air fuel mixer for a high pressure ratio aero-derivative gas turbine engine which uniformly mixes fuel and air without incurring backflow from the combustor.
Still another objective of the present invention is to inject fuel into an air fuel mixer in such a manner as to maximize mixing therein.
These objectives and other features of the present invention will become more readily apparent upon reference to the following description when taken in conjunction with the following drawing.